The present invention relates generally to gas turbine engines, and, more specifically, to compressors therein.
In a gas turbine engine air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases. The combustion gases are discharged through turbine stages which extract energy therefrom for powering the compressor, and producing output power for use in driving a fan in an exemplary turbofan aircraft engine application.
A multistage axial compressor includes cooperating rows of stator vanes and rotor blades which decrease in size to pressurize air in stages. The compressor vanes and blades have corresponding airfoils which typically vary in configuration as their size decreases from stage to stage for maximizing performance of the compressor. Compressor performance includes, for example, efficiency of compression, flow capability, and stall margin, which are all affected by the configuration of the vanes and blades.
More specifically, the flow or pressure distribution of the air as it is being compressed through the stator vanes and rotor blades is a complex three dimensional flow field varying circumferentially around the compressor, radially along the span of the vane and blade airfoils, and axially along the circumferentially opposite pressure and suction sides of the airfoils
The airfoil pressure side is a generally concave surface cooperating with the opposite suction side, which is a generally convex surface, for efficiently pressurizing the air as it flows between blades in the axial downstream direction between the leading and trailing edges thereof. The pressure distribution of the air undergoing compression varies from the radially inner root of the airfoil to the radially outer tip of the airfoil which is spaced closely adjacent to a surrounding compressor casing to provide a suitable radial gap or clearance therewith.
The airfoil, itself, may be supported from the compressor rotor in any suitable manner such as being formed integrally therewith in a unitary blisk configuration, or each rotor airfoil may have an integral platform and dovetail for mounting the compressor blade in a corresponding dovetail slot formed in the perimeter of the compressor rotor.
A significant feature affecting compressor performance is the radial clearance provided between the airfoil tips and surrounding casing. The clearance should be as small as possible to minimize undesirable flow losses therethrough, but must be sufficiently large for accommodating transient operation of the compressor which may occasionally lead to tip rubs. In a tip rub, material is removed from the airfoil tip and may accumulate on the inner surface of the casing. The shortened tip increases the clearance with the casing which decreases compressor performance, which is further affected by any accumulation of rub material on the casing which disrupts the smooth flow of air therealong.
Nevertheless, commercial experience of multistage axial compressors in aircraft turbofan engines confirms long useful lives for the compressor rotor blades and continued high performance of the compressor. However, the occasional compressor blade tip rubs increase tip clearances and decrease compressor performance over the useful blade lifetime. The loss in compressor performance due to tip rubs further reduces performance of the engine since the pressurized air is used in the combustion process, and energy is extracted from the combustion gases in the turbines.
Accordingly, it is desired to provide a compressor rotor airfoil having improved aerodynamic efficiency notwithstanding increased blade tip clearances due to tip rubs.